With reference to FIG. 1, a ducted fan gas turbine engine generally indicated at 10 has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, and intermediate pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle 23.
The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate pressure compressor 13 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 16, 17, 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.
The performance of gas turbine engines, whether measured in terms of efficiency or specific output, is improved by increasing the turbine gas temperature. It is therefore desirable to operate the turbines at the highest possible temperatures. For any engine cycle compression ratio or bypass ratio, increasing the turbine entry gas temperature produces more specific thrust (e.g. engine thrust per unit of air mass flow). However as turbine entry temperatures increase, the life of an un-cooled turbine falls, necessitating the development of better materials and the introduction of internal air cooling.
In modern engines, the high-pressure turbine gas temperatures are hotter than the melting point of the material of the blades and vanes, necessitating internal air cooling of these airfoil components. During its passage through the engine, the mean temperature of the gas stream decreases as power is extracted. Therefore, the need to cool the static and rotary parts of the engine structure decreases as the gas moves from the high-pressure stage(s), through the intermediate-pressure and low-pressure stages, and towards the exit nozzle.
FIG. 2 shows an isometric view of a typical single stage cooled turbine. Cooling air flows are indicated by arrows.
Internal convection and external films are the prime methods of cooling the gas path components—airfoils, platforms, shrouds and shroud segments etc. High-pressure turbine nozzle guide vanes 31 (NGVs) consume the greatest amount of cooling air on high temperature engines. High-pressure blades 32 typically use about half of the NGV flow. The intermediate-pressure and low-pressure stages downstream of the HP turbine use progressively less cooling air.
The high-pressure turbine airfoils are cooled by using high pressure air from the compressor that has by-passed the combustor and is therefore relatively cool compared to the gas temperature. Typical cooling air temperatures are between 800 and 1000 K, while gas temperatures can be in excess of 2100 K.
The cooling air from the compressor that is used to cool the hot turbine components is not used fully to extract work from the turbine. Therefore, as extracting coolant flow has an adverse effect on the engine operating efficiency, it is important to use the cooling air effectively.
Ever increasing gas temperature levels combined with a drive towards flatter combustion radial profiles, in the interests of reduced combustor emissions, have resulted in an increase in local gas temperature experienced by the extremities of the blades and vanes, and the working gas annulus endwalls.
A turbine blade or vane has a longitudinally extending aerofoil portion with facing suction side and pressure side walls. The aerofoil portion extends across the working gas annulus, with the longitudinal direction of the aerofoil portion being along a radial direction of the engine. FIG. 3 shows a longitudinal cross-section through a high-pressure turbine blade. A multi-pass cooling passage 33 is fed cooling air by a feed passage 34 at the root of the blade. Cooling air eventually leaves the multi-pass cooling passage through exit holes at the tip 35 and the trailing edge 36 of the blade. Some of the cooling air, however, can leave the multi-pass cooling passage through effusion holes (not shown) formed in the suction side and pressure side walls. The block arrows in FIG. 3 show the general direction of cooling air flow.
The (triple) multi-pass cooling passage 33 is formed by two divider walls 37 which interconnect the facing suction side and pressure side walls of the aerofoil portion to form three longitudinally extending, side-by-side passage portions 38. Other aerofoil portions can have more or fewer divider walls and passage portions. The passage portions are connected in series fluid flow relationship by respective bends 39 which are formed by the joined ends of neighbouring passage portions. The cooling air thus enters the multi-pass cooling passage at the passage portion at the leading edge of the aerofoil portion and flows through each passage portion in turn to eventually leave from the passage portion at the trailing edge. Trip strip 40 and pedestal 41 heat transfer augmentation devices in the passage portions enhance heat transfer between the cooling air and the metal.
The internal convection is achieved by two mechanisms, firstly through augmented channel flow, or impingement cooling inside the cooling passage, and secondly by the internal convection inside the film cooling holes.
The complicated internal structure of the aerofoil portion is generally formed by an investment casting procedure. Thus the mould for the aerofoil portion has a core structure which is a “negative” of the ultimate internal structure of the aerofoil portion. In particular, the mould has passage features corresponding to the longitudinally extending passage portions 38.
In order to maintain acceptable component lives in particularly the high pressure turbine rotor blade, more effective cooling schemes have been adopted, such as impingement leading edge cooling arrangements and cyclonic or forced vortex cooling, specifically in the vicinity of the aerofoil leading edge, where the external heat load is at its greatest.
US2006/0280607 proposes an aerofoil arrangement in which a first cooling passage has a tangential inlet flow from a neighbouring cooling passage. This develops a vortex in the first passage due to the momentum and direction of the flow. The vortex can be arranged to rotate in either a clockwise or anticlockwise direction by changing the location of the tangential feed channel. The strength of the vortex is a function of the flow rate, which is dependent on the pressure ratio between the passages and the flow area, and also the physical geometry (angle and location) of the tangential feed passage. However, the strength of the vortex reduces as flow is extracted up the span of the aerofoil.
It is possible to manufacture both this arrangement using conventional ceramic cores produced by multi-pull core dies or with soluble core technology.